U.S. patent number 7,195,458 [Application Number 10/884,440] was granted by the patent office on 2007-03-27 for impingement cooling system for a turbine blade.
This patent grant is currently assigned to Siemens Power Generation, Inc.. Invention is credited to George Liang.
United States Patent |
7,195,458 |
Liang |
March 27, 2007 |
Impingement cooling system for a turbine blade
Abstract
A turbine blade for a turbine engine having a leading edge
cooling system formed from a suction side cooling channel and a
pressure side cooling channel. Cooling fluids flow into the leading
edge cooling channels through impingement orifices that meter
cooling fluid flow. The cooling fluids may form vortices in the
cooling channels before being released from the turbine blade
through gill holes. The cooling fluids then form a boundary layer
of film cooling fluids on an outer surface of the turbine
blade.
Inventors: |
Liang; George (Palm City,
FL) |
Assignee: |
Siemens Power Generation, Inc.
(Orlando, FL)
|
Family
ID: |
35514094 |
Appl.
No.: |
10/884,440 |
Filed: |
July 2, 2004 |
Prior Publication Data
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|
|
|
Document
Identifier |
Publication Date |
|
US 20060002795 A1 |
Jan 5, 2006 |
|
Current U.S.
Class: |
416/97R |
Current CPC
Class: |
F01D
5/187 (20130101); F05D 2260/201 (20130101); F05D
2260/202 (20130101); F05D 2260/22141 (20130101) |
Current International
Class: |
F01D
5/18 (20060101) |
Field of
Search: |
;415/115-116
;416/96R,96A,97R |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
Primary Examiner: Verdier; Christopher
Claims
I claim:
1. A turbine blade, comprising: a generally elongated blade having
a leading edge, a trailing edge, and a tip at a first end, a root
coupled to the blade at an end generally opposite the first end for
supporting the blade, at least one cavity forming a cooling system
in the blade, and an outer wall defining the at least one cavity
forming the cooling system; wherein the cooling system comprises a
leading edge cooling channel formed from pressure side cooling
channels extending radially within the elongated blade and suction
side cooling channels extending radially within the elongated blade
and separated from the pressure side cooling channels by a rib;
wherein the pressure side cooling channels each include at least
one impingement orifice providing a fluid pathway between the
pressure side cooling channels and other portions of the cooling
system; and wherein the suction side cooling channels each include
at least one impingement orifice providing a fluid pathway between
the suction side cooling channels and other portions of the cooling
system.
2. The turbine blade of claim 1, further comprising at least one
gill hole in the outer wall providing a fluid pathway between at
least one of the suction side cooling channels and an outer surface
of the turbine blade and positioned to exhaust a cooling fluid in a
general downstream direction.
3. The turbine blade of claim 1, further comprising at least one
gill hole in the outer wall providing a fluid pathway between at
least one of the pressure side cooling channels and an outer
surface of the turbine blade and positioned to exhaust a cooling
fluid in a general downstream direction.
4. The turbine blade of claim 1, wherein the at least one
impingement orifice in the suction side cooling channels comprises
a filleted inlet and a filleted outlet.
5. The turbine blade of claim 1, wherein the at least one
impingement orifice in the pressure side cooling channels comprises
a filleted inlet and a filleted outlet.
6. The turbine blade of claim 1, wherein the at least one
impingement orifice in the pressure side cooling channels is
positioned proximate to the rib separating the pressure side
cooling channels from the suction side cooling channels to pass
cooling fluids along the rib to form a vortex.
7. The turbine blade of claim 1, wherein the at least one
impingement orifice in the suction side cooling channels is
positioned proximate to the rib separating the pressure side
cooling channels from the suction side cooling channel to pass
cooling fluids along the rib to form a vortex.
8. The turbine blade of claim 1, wherein the suction side cooling
channels are offset from the pressure side cooling channels in a
spanwise direction.
9. The turbine blade of claim 1, wherein there are five suction
side cooling channels and three pressure side cooling channels.
10. The turbine blade of claim 1, wherein the at-least pressure
side cooling channel channels comprise a plurality of channels
aligned in a spanwise direction along the leading edge.
11. A turbine blade, comprising: a generally elongated blade having
a leading edge, a trailing edge, and a tip at a first end, a root
coupled to the blade at an end generally opposite the first end for
supporting the blade, at least one cavity forming a cooling system
in the blade, and an outer wall defining the at least one cavity
forming the cooling system; wherein the cooling system comprises a
leading edge cooling channel formed from a plurality of pressure
side cooling channels extending radially within the elongated blade
and a plurality of suction side cooling channels extending radially
within the elongated blade, offset spanwise relative to the
pressure side cooling channels, and separated from the pressure
side cooling channel by a rib; wherein the pressure side cooling
channels include at least one impingement orifice providing a fluid
pathway between the pressure side cooling channels and other
portions of the cooling system; and wherein the suction side
cooling channels include at least one impingement orifice providing
a fluid pathway between the suction side cooling channels and other
portions of the cooling system.
12. The turbine blade of claim 11, further comprising at least one
gill hole in the outer wall providing a fluid pathway between at
least one of the suction side cooling channels and an outer surface
of the turbine blade.
13. The turbine blade of claim 12, further comprising at least one
gill hole in the outer wall providing a fluid pathway between at
least one of the pressure side cooling channels and an outer
surface of the turbine blade, wherein the gill holes in the suction
side cooling channels and the pressure side cooling channels are
positioned to exhaust a cooling fluid in a general downstream
direction.
14. The turbine blade of claim 11, wherein the at least one
impingement orifice in the suction side cooling channels comprise a
filleted inlet and a filleted outlet.
15. The turbine blade of claim 11, wherein the at least one
impingement orifice in the pressure side cooling channels comprise
a filleted inlet and a filleted outlet.
16. The turbine blade of claim 11, wherein the at least one
impingement orifice in the pressure side cooling channel is
positioned proximate to the rib separating the pressure side
cooling channel from the suction side cooling channel to pass
cooling fluids along the rib to form a vortex, and the at least one
impingement orifice in the suction side cooling channel is
positioned proximate to the rib separating the pressure side
cooling channel from the suction side cooling channel to pass
cooling fluids along the rib to form a vortex.
17. The turbine blade of claim 11, wherein there are five suction
side cooling channels and three pressure side cooling channels.
Description
FIELD OF THE INVENTION
This invention is directed generally to turbine blades, and more
particularly to hollow turbine blades having internal cooling
channels for passing cooling fluids, such as air, through the
cooling channels to cool the blades.
BACKGROUND
Typically, gas turbine engines include a compressor for compressing
air, a combustor for mixing the compressed air with fuel and
igniting the mixture, and a turbine blade assembly for producing
power. Combustors often operate at high temperatures that may
exceed 2,500 degrees Fahrenheit. Typical turbine combustor
configurations expose turbine blade assemblies to these high
temperatures. As a result, turbine blades must be made of materials
capable of withstanding such high temperatures. In addition,
turbine blades often contain cooling systems for prolonging the
life of the blades and reducing the likelihood of failure as a
result of excessive temperatures.
Typically, turbine blades are formed from a root portion and a
platform at one end and an elongated portion forming a blade that
extends outwardly from the platform. The blade is ordinarily
composed of a tip opposite the root section, a leading edge, and a
trailing edge. The inner aspects of most turbine blades typically
contain an intricate maze of cooling channels forming a cooling
system. The cooling channels in the blades receive air from the
compressor of the turbine engine and pass the air through the
blade. The cooling channels often include multiple flow paths that
are designed to maintain all aspects of the turbine blade at a
relatively uniform temperature. However, centrifugal forces and air
flow at boundary layers often prevent some areas of the turbine
blade from being adequately cooled, which results in the formation
of localized hot spots. Localized hot spots, depending on their
location, can reduce the useful life of a turbine blade and can
damage a turbine blade to an extent necessitating replacement of
the blade.
Conventional turbine blades often include a plurality of holes in
the leading edges that form showerheads for exhausting cooling
fluids from the internal cooling systems to be used as film cooling
fluids on the outer surfaces of the turbine blades. Often times,
the cooling fluids flowing through these holes are not regulated.
Instead, cooling fluids are often passed through the showerhead at
too high of a flow rate, which create turbulence in boundary layers
of cooling fluids at the outer surfaces of the turbine blades. This
turbulence reduces the effectiveness of downstream film cooling. In
addition, the cooling fluids are often discharged at dissimilar
pressures, which further reduces the downstream film cooling
effectiveness. While these conventional systems reduce the
temperature of leading edges of turbine blades, a need exist for an
improved leading edge cooling system capable of operating more
efficiently.
SUMMARY OF THE INVENTION
This invention relates to a turbine blade cooling system of a
turbine engine. In particular, the cooling system includes a
multiple channel leading edge cooling system for removing heat from
the leading edge of a turbine blade. The turbine blade may be
generally elongated and have a leading edge, a trailing edge, a tip
at a first end, a root coupled to the blade at an end opposite the
first end for coupling the blade to the disc, and at least one
cavity forming at least a portion of the cooling system. The
cooling system may be formed from a leading edge cooling channel
formed from a pressure side cooling channel extending radially
within the elongated blade and a suction side cooling channel
extending radially within the elongated blade and separated from
the pressure side cooling channel by a rib. The pressure side
cooling channel may include at least one impingement orifice
providing a fluid pathway between the pressure side cooling channel
and other portions of the cooling system. In addition, the suction
side cooling channel may include at least one impingement orifice
providing a fluid pathway between the suction side cooling channel
and other portions of the cooling system. The impingement orifices
may be offset within the cooling channels such that cooling fluids
are directed to flow generally along the rib separating the suction
side and pressure side cooling channels to form vortices in the
cooling channels. The impingement orifices may include filleted
inlets and filleted outlets as well.
In at least one embodiment, the leading edge cooling channel may be
formed from a plurality of cooling channels that regulate the flow
of cooling fluids through the cooling system. For instance, there
may be, but is not limited to, about three pressure side cooling
channels and about five suction side cooling channels. The cooling
channels may be offset from each other in the spanwise direction to
increase convection in the channels. In other embodiments, the
suction side and pressure side cooling channels may be aligned in
the spanwise direction.
The cooling system may also include one or more gill holes in the
outer wall providing a fluid pathway between the suction side
cooling channel and an outer surface of the turbine blade. The gill
holes may be located in the suction side cooling channel or the
pressure side cooling channel, or both. The gill holes may be
positioned in the cooling channels such that cooling fluids
exhausted through the gill holes are not directed directly into
oncoming combustion gases. Rather, the gill holes may be positioned
in the outer wall such that cooling fluids exhausted from the gill
holes are directed generally downstream with the flow of combustion
gases.
In operation, cooling fluids, which may be air and other gases, are
passed into the cooling system through the root of a blade from a
compressor or other source. At least a portion of the cooling
fluids flow through the impingement orifices into the leading edge
cooling channels. For instance, the cooling fluids may flow through
the impingement orifices and form vortices in the cooling channels.
As the cooling fluids spin within the cooling channels and contact
the walls forming the cooling channels, the cooling fluids increase
in temperature. The cooling fluids are exhausted from the cooling
channels through the gill holes. Because of the angle of the gill
holes, the cooling fluids exhausted by the gill holes are not
dispersed into the main flow of combustion gases. Rather, the
cooling fluids form a layer of film cooling fluids at an outer
surface of the turbine blade.
An advantage of this invention is that the impingement orifices
meter the flow of cooling fluids that enter the leading edge
cooling channel, thereby controlling the temperature of the leading
edge.
Another advantage of this invention is that the impingement
orifices limit the flow of cooling fluids from the gill holes and
thereby limit cooling fluid penetration into the flow of combustion
gases, yielding a desirable coolant sub-boundary layer at the outer
surface of the turbine blade.
Yet another advantage of this invention is that the position of the
impingement holes create vortices in the suction side and pressure
side cooling channels that increase convection in these areas and
increase heat removal from the outer wall proximate to the
stagnation region.
Another advantage of this invention is that the compartmentalized
leading edge cooling channel maximizes usage of the cooling fluid
for a particular turbine blade inlet gas temperature and pressure
profile.
Still another advantage of this invention is that by offsetting the
pressure side cooling channels relative to the suction side cooling
channels the amount of heat reduction is increased.
These and other embodiments are described in more detail below.
BRIEF DESCRIPTION OF THE DRAWINGS
The accompanying drawings, which are incorporated in and form a
part of the specification, illustrate embodiments of the presently
disclosed invention and, together with the description, disclose
the principles of the invention.
FIG. 1 is a perspective view of a turbine blade containing a
cooling system of this invention.
FIG. 2 is a partial cross-sectional view of the leading edge
cooling system of this invention taken along section line 2--2 in
FIG. 1.
FIG. 3 is a cross-sectional view of the turbine blade of FIG. 1
taken along section line 3--3 showing the pressure side cooling
channels.
FIG. 4 is cross-sectional view of the turbine blade of FIG. 1 taken
along section line 4--4 showing the suction side cooling
channels.
FIG. 5 is partial cross-sectional view of an alternative embodiment
of the leading edge cooling channels taken along section line 2--2
in FIG. 1.
DETAILED DESCRIPTION OF THE INVENTION
As shown in FIGS. 1 5, this invention is directed to a turbine
blade cooling system 10 for turbine blades 12 used in turbine
engines. In particular, turbine blade cooling system 10 is directed
to a cooling system 10 located in a cavity 14, as shown in FIGS. 3
and 4, positioned between outer walls 22. Outer walls 22 form a
housing 24 of the turbine blade 12. As shown in FIG. 1, the turbine
blade 12 may be formed from a root 16 having a platform 18 and a
generally elongated blade 20 coupled to the root 16 at the platform
18. The turbine blade may also include a tip 36 generally opposite
the root 16 and the platform 18. Blade 20 may have an outer wall 22
adapted for use, for example, in a first stage of an axial flow
turbine engine. Outer wall 22 may have a generally concave shaped
portion forming pressure side 26 and may have a generally convex
shaped portion forming suction side 28.
The cavity 14, as shown in FIGS. 3 and 4, may be positioned in
inner aspects of the blade 20 for directing one or more gases,
which may include air received from a compressor (not shown),
through the blade 20 and out one or more orifices 34 in the blade
20. As shown in FIGS. 3 and 4, the orifices 34 may be positioned in
a leading edge 38, a trailing edge 40, the pressure side 26, and
the suction side 28 to provide film cooling. The orifices 34
provide a pathway from the cavity 14 through the outer wall 22.
As shown in FIG. 2, the cavity 14 forming the cooling system 10 may
include one or more leading edge cooling cavities 42. The leading
edge cooling cavity 42 may be formed from a suction side cooling
channel 44 extending radially within the blade 20 and a pressure
side cooling channel 46 extending radially within the blade 20. The
suction and pressure side cooling channels 44, 46 may be separated
by a rib 47. The suction and pressure side cooling channels 44, 46
may extend from the root 16 to the tip 36, or in other embodiments,
may extend radially along only a portion of the leading edge 38. In
at least one embodiment, as shown in FIG. 4, the suction side
cooling channel 44 may be formed from a plurality of channels. For
instance, the cooling system 10 may include, but is not limited to,
five suction side cooling channels 44. The pressure side cooling
channel 46 may also be formed from a plurality of channels. For
instance, the cooling system 10 may include, but is not limited to,
three pressure side cooling channels 46. The suction and pressure
side cooling channels 44, 46 may be aligned radially along the
leading edge 38. In alternative embodiments, the suction and
pressure side cooling channels 44, 46 may be offset radially in the
spanwise direction as shown in FIGS. 3 and 4. Offsetting the
suction and pressure side cooling channels 44, 46 increases the
ability of the channels 44, 46 to dissipate heat from the blade 20
to the cooling fluid flowing through the cooling system 10.
As shown in FIGS. 2 4, the cooling system 10 may include one or
more impingement orifices 48 providing a fluid pathway between the
suction side cooling channel 44 and other portions of the cooling
system 10. The impingement orifice 48 may extend through a rib 60
separating the leading edge cooling cavity 42 from other aspects of
the cavity 14. There may exist one impingement orifice or a
plurality of impingement orifices along the length of the suction
side cooling channel 44. The impingement orifice 44 may include a
filleted inlet 50 and a filleted outlet 52. Similarly, the cooling
system 10 may include one or more impingement orifices 54 providing
a fluid pathway between the pressure side cooling channel 46 and
other portions of the cooling system 10. There may exist one
impingement orifice or a plurality of impingement orifices 54 along
the length of the pressure side cooling channel 46. The impingement
orifice 54 may include a filleted inlet 56 and a filleted outlet
58.
In at least one embodiment, as shown in FIG. 5, the impingement
orifice 48 may be positioned such that the outlet 52 is in close
proximity with the rib 47 and the fluid flowing through the
impingement orifice 48 is directed to flow generally along the rib
47 and to form a vortex in the suction side cooling channel 44.
Formation of the vortex may increase the ability of the impingement
orifice 48 to remove heat from the blade 20, and more particularly,
reduces the temperature of the outer wall 22 proximate to the
stagnation point 66. Similarly, the impingement orifice 54 may be
positioned such that the outlet 58 is in close proximity with the
rib 47 and the fluid flowing through the impingement orifice 54 is
directed to flow generally along the rib 47 and to form a vortex in
the pressure side cooling channel 46.
The cooling system 10 may also include one or more gill holes 62 in
the outer wall 22 providing a fluid pathway between the suction
side cooling channel 44 and an outer surface 64 of the blade 20.
The gill holes 62 may also provide a fluid pathway between the
pressure side cooling channel 46 and the outer surface 64 of the
blade 20. The gill hole 62 may be positioned such that the fluids
exhausted from the suction side cooling channel 44 are not directed
directly into the oncoming combustion gases. Rather, the gill holes
62 are positioned to exhaust cooling fluids from the cooling system
10 generally in the downstream direction of flow of the combustion
gases past the blade 20.
During operation, cooling fluids enter the cooling system 10
through the root 16 as typically supplied from a compressor. The
cooling fluids flow through various aspects of the cooling system
and are exhausted through orifices 34. At least a portion of the
cooling fluids is passed into the leading edge cooling cavity 42
through the impingement orifices 48 and 54. As the cooling fluids
enter the suction and pressure side cooling channels 44, 46, the
cooling fluids pass along the rib 47 and form vortices in the
channels 44, 46. The fluids accept heat from the surface of the rib
47, rib 60, and the outer wall 22. The cooling fluids are exhausted
through the gill holes 62 in the outer wall 22 and function as film
cooling fluids on the outer surface 64 of the outer wall 22.
The foregoing is provided for purposes of illustrating, explaining,
and describing embodiments of this invention. Modifications and
adaptations to these embodiments will be apparent to those skilled
in the art and may be made without departing from the scope or
spirit of this invention.
* * * * *